# Pressure Ratios for AeroEngines

Pressure Ratios for AeroEngines GAS TURBINES Pressure Ratios for reasons, it is desired that the low pressure system should do less work, the high pressure turbine Aero-Engines will be overworked, or overweight (i.e. have two stages). A Note on the Factors Limiting the TABL E 1 Turbojet s Choice of Pressure Ratio Number of turbin e stages 1 2 3 By J. M. Stephenson 4 1 10-2 18-7 ' C Division of turbine stages, 1 1:1 1:2 low pressure : high pressure 2:1 Discussion T is usually supposed that the designer of a gas Two-spool By-pass turbine engine has considerable freedom in In the by-pass engine the relation between the choosing the pressures, temperatures and other compressor and turbine pressure ratios depends features. Indeed this apparent flexibility has been on the by-pass ratio, or the ratio between the mass claimed as an inherent advantage of the gas of air tapped off the low pressure compressor to turbine, which enables it to be tailor-made for that passing through the high pressure compressor any prescribed duty. and the turbines. For illustration this will be This may be true for ground installations, but assumed equal, to 0-5, though any other value in aircraft engines there is a serious restriction on could be taken. An analysis similar to that for the the choice of pressure ratio, for, to keep down the jet engine gives the curve B in FIG. 1. The possible weight, each single compressor and turbine stage engines with the given turbine stage pressure ratio must be designed to handle as much work as can be picked straight off this curve, and are listed possible. It is assumed in this note that there is, in TADLE II. for aerodynamic reasons, a unique upper limit to the pressure ratio that can be attained effi­ TABL E II ciently by a turbine stage; and that the weight of Two-spool By-pass of Ratio 0-5 a turbine stage is so nearly independent of its pressure ratio that it is always designed to have Number of turbine stages 2 3 ■4 the limiting value, which is here set arbitrarily at 7' 8 12-5 18-1 1-80. It is shown that this assumption restricts >c the possible engine layouts.whether jet, by-pass, or Division of turbii e stage*, 2:1 3:1 turboprop, to quite a small number. If the present low pressure : high pressure 1:1 1:2 2:2 REFERENCE TO LITERATURE 1:3 assumption is close to the truth, future engines are J. M. Stephenson: A Solution of the Surging Problem in Axial- likely to crystallize into these few types, though Flow Compressors, Journal of the Aeronautical Sciences, January it is true that some existing engines fall outside 1952, p. 67. Turboprop the pattern, by using low turbine pressure ratios. If it is found possible to use turbine stage In this case practically all theexcess power from Turbojet pressure ratios much greater than 1-8, only the the gas generator is delivered to the propeller, numerical values are affected in the argument that and the turbine and compressor pressure ratios The essential feature of the turbojet engine is follows. are nearly equal, or that the gas generator is self-contained; i.e. that Similar considerations do not have to be ap­ the turbine just drives the compressor. Hence, r =0-93r t c plied to the compressor stages, which have much with the usual notation, where the odd 7 per cent is assumed to be lost in smaller pressure ratios than turbines, so that any C AT =C \T the combustion system, or delivered to the jet vc c rl t ratio can be approached quite closely with a pipe. This is shown as curve P in FIG 1 and the where C„ cc y/(y—1) whole number of stages. possible engines arc set out in TABLE HI. It is well known that a thermally efficient Also i7ir/r=r<»-1)/)'-l for the com­ c c c c engine requires as high a pressure ratio as possible pressor, TABL E III irrespective of the type of propulsion used. The and &Tl-qTt=\-\lriY-^iY for the tur­ t t t Turboprop s difficulty of making a single shaft compressor run bine. stably, especially during acceleration, when the Therefore, substituting y=7/ 5 for the compres­ Number of turbin e stages 3 4 5 design pressure ratio is more than about six, is sor, and y=4/ 3 for the turbine, putting both 6-3 II I 20-2 discussed in Ref. 1. The most compact and simple 'c efficiencies equal to 0-90, and the entry tempera­ solution to the stability problem is to use a 'two- ture ratio TJT =4, the relation between the Division of turbine stages, 3 4:1 3:2 spool' gas generator, where both the compressor high pressure : low pressure pressure ratios is 1:2 4:1 and turbine are split and run on independent (] / ,) 1/ 4 = l-27-0-27(r.)8'7 coaxial shafts. With this arrangement, pressure r This is plotted in FIG 1 as curve J. It may be ratios up to about 20 can be achieved; beyond Note that the 1: 2 arrangement implies no seen that if the turbine stage pressure ratio is compression on the low pressure shaft, or a 'free this it is probable that three shafts would have to 1 • 8, the only possible jet engines are those shown power turbine'. This effect would also occur with be used, but this type will not be discussed here. in TABLE i. Note that although the turbine pres­ a 2 : 3 arrangement on an engine of pressure The requirement that the turbine be split re­ sure ratios are equal in the 10 : 1 engine, the low ratio 20, but it may be assumed that the com­ stricts the possible division of work between the pressor would be unstable. shafts, in some cases fixing the value uniquely. pressure system does more work. If, for other December 1953 371 http://www.deepdyve.com/assets/images/DeepDyve-Logo-lg.png Aircraft Engineering and Aerospace Technology Emerald Publishing

# Pressure Ratios for AeroEngines

, Volume 25 (12): 1 – Dec 1, 1953
1 page

/lp/emerald-publishing/pressure-ratios-for-aeroengines-P1og7w4JZf

# References

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Publisher
Emerald Publishing
ISSN
0002-2667
DOI
10.1108/eb032365
Publisher site
See Article on Publisher Site

### Abstract

GAS TURBINES Pressure Ratios for reasons, it is desired that the low pressure system should do less work, the high pressure turbine Aero-Engines will be overworked, or overweight (i.e. have two stages). A Note on the Factors Limiting the TABL E 1 Turbojet s Choice of Pressure Ratio Number of turbin e stages 1 2 3 By J. M. Stephenson 4 1 10-2 18-7 ' C Division of turbine stages, 1 1:1 1:2 low pressure : high pressure 2:1 Discussion T is usually supposed that the designer of a gas Two-spool By-pass turbine engine has considerable freedom in In the by-pass engine the relation between the choosing the pressures, temperatures and other compressor and turbine pressure ratios depends features. Indeed this apparent flexibility has been on the by-pass ratio, or the ratio between the mass claimed as an inherent advantage of the gas of air tapped off the low pressure compressor to turbine, which enables it to be tailor-made for that passing through the high pressure compressor any prescribed duty. and the turbines. For illustration this will be This may be true for ground installations, but assumed equal, to 0-5, though any other value in aircraft engines there is a serious restriction on could be taken. An analysis similar to that for the the choice of pressure ratio, for, to keep down the jet engine gives the curve B in FIG. 1. The possible weight, each single compressor and turbine stage engines with the given turbine stage pressure ratio must be designed to handle as much work as can be picked straight off this curve, and are listed possible. It is assumed in this note that there is, in TADLE II. for aerodynamic reasons, a unique upper limit to the pressure ratio that can be attained effi­ TABL E II ciently by a turbine stage; and that the weight of Two-spool By-pass of Ratio 0-5 a turbine stage is so nearly independent of its pressure ratio that it is always designed to have Number of turbine stages 2 3 ■4 the limiting value, which is here set arbitrarily at 7' 8 12-5 18-1 1-80. It is shown that this assumption restricts >c the possible engine layouts.whether jet, by-pass, or Division of turbii e stage*, 2:1 3:1 turboprop, to quite a small number. If the present low pressure : high pressure 1:1 1:2 2:2 REFERENCE TO LITERATURE 1:3 assumption is close to the truth, future engines are J. M. Stephenson: A Solution of the Surging Problem in Axial- likely to crystallize into these few types, though Flow Compressors, Journal of the Aeronautical Sciences, January it is true that some existing engines fall outside 1952, p. 67. Turboprop the pattern, by using low turbine pressure ratios. If it is found possible to use turbine stage In this case practically all theexcess power from Turbojet pressure ratios much greater than 1-8, only the the gas generator is delivered to the propeller, numerical values are affected in the argument that and the turbine and compressor pressure ratios The essential feature of the turbojet engine is follows. are nearly equal, or that the gas generator is self-contained; i.e. that Similar considerations do not have to be ap­ the turbine just drives the compressor. Hence, r =0-93r t c plied to the compressor stages, which have much with the usual notation, where the odd 7 per cent is assumed to be lost in smaller pressure ratios than turbines, so that any C AT =C \T the combustion system, or delivered to the jet vc c rl t ratio can be approached quite closely with a pipe. This is shown as curve P in FIG 1 and the where C„ cc y/(y—1) whole number of stages. possible engines arc set out in TABLE HI. It is well known that a thermally efficient Also i7ir/r=r<»-1)/)'-l for the com­ c c c c engine requires as high a pressure ratio as possible pressor, TABL E III irrespective of the type of propulsion used. The and &Tl-qTt=\-\lriY-^iY for the tur­ t t t Turboprop s difficulty of making a single shaft compressor run bine. stably, especially during acceleration, when the Therefore, substituting y=7/ 5 for the compres­ Number of turbin e stages 3 4 5 design pressure ratio is more than about six, is sor, and y=4/ 3 for the turbine, putting both 6-3 II I 20-2 discussed in Ref. 1. The most compact and simple 'c efficiencies equal to 0-90, and the entry tempera­ solution to the stability problem is to use a 'two- ture ratio TJT =4, the relation between the Division of turbine stages, 3 4:1 3:2 spool' gas generator, where both the compressor high pressure : low pressure pressure ratios is 1:2 4:1 and turbine are split and run on independent (] / ,) 1/ 4 = l-27-0-27(r.)8'7 coaxial shafts. With this arrangement, pressure r This is plotted in FIG 1 as curve J. It may be ratios up to about 20 can be achieved; beyond Note that the 1: 2 arrangement implies no seen that if the turbine stage pressure ratio is compression on the low pressure shaft, or a 'free this it is probable that three shafts would have to 1 • 8, the only possible jet engines are those shown power turbine'. This effect would also occur with be used, but this type will not be discussed here. in TABLE i. Note that although the turbine pres­ a 2 : 3 arrangement on an engine of pressure The requirement that the turbine be split re­ sure ratios are equal in the 10 : 1 engine, the low ratio 20, but it may be assumed that the com­ stricts the possible division of work between the pressor would be unstable. shafts, in some cases fixing the value uniquely. pressure system does more work. If, for other December 1953 371

### Journal

Aircraft Engineering and Aerospace TechnologyEmerald Publishing

Published: Dec 1, 1953